DOCSLIB.ORG

Expander Demonstrator

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text

Load more

FLPP_134.qxd 6/5/08 12:37 PM Page 34 Expander Demonstrator Jérôme Breteau & Jean-Noël Caruana Future Launcher Preparatory Programme, Directorate of Launchers, ESA HQ, Paris, France he technical difficulties encountered by other spacefaring countries in similar T ‘expander-cycle’ engine projects show how demanding it is to master this kind of technology. So the achievements of ESA’s Future Launchers Preparatory Programme (FLPP) Expander Demonstrator Project, including several European ‘firsts’, are essential contributions to the development of future cryogenic upper-stage engines. Introduction In launch vehicles, one of the key enabling technologies is the propulsion system, but this is complicated to acquire and takes a long time in development. This is especially true for upper-stage engines, where use of cryogenic propellants such as liquid hydrogen and oxygen and reignition capability are essential in order to reach high-energy orbits with heavy payloads. Upper-stage engines also operate in specific conditions (vacuum, micro- gravity) that are difficult to reproduce on Earth, and involve significant development risks that have to be mitigated. esa bulletin 134 - may 2008 35 FLPP_134.qxd 6/5/08 12:37 PM Page 34 Expander Demonstrator Jérôme Breteau & Jean-Noël Caruana Future Launcher Preparatory Programme, Directorate of Launchers, ESA HQ, Paris, France he technical difficulties encountered by other spacefaring countries in similar T ‘expander-cycle’ engine projects show how demanding it is to master this kind of technology. So the achievements of ESA’s Future Launchers Preparatory Programme (FLPP) Expander Demonstrator Project, including several European ‘firsts’, are essential contributions to the development of future cryogenic upper-stage engines. Introduction In launch vehicles, one of the key enabling technologies is the propulsion system, but this is complicated to acquire and takes a long time in development. This is especially true for upper-stage engines, where use of cryogenic propellants such as liquid hydrogen and oxygen and reignition capability are essential in order to reach high-energy orbits with heavy payloads. Upper-stage engines also operate in specific conditions (vacuum, micro- gravity) that are difficult to reproduce on Earth, and involve significant development risks that have to be mitigated. esa bulletin 134 - may 2008 35 FLPP_134.qxd 6/5/08 12:37 PM Page 36 Launchers Expander Demonstrator In 1998 ESA, CNES and Arianespace project called the ‘Cryogenic Reignitable (TRL) of the engine and its decided to develop an enhanced Upper Stage Engine – Expander components, currently estimated cryogenic upper-stage for the Ariane-5 Demonstrator’. between 3 and 5, depending on the launcher in order to respond to the subjects under study, must be raised to rapid evolution of the global Project Objectives level 6 (prototype test in a relevant commercial market towards more heavy The main aim of the FLPP Expander environment), in order to properly assess payloads. In recognition of the quick Demonstrator Project is to supply the risks, cost and duration of a further emergence of this new commercial need elements allowing a sound, informed development phase up to qualification and the comparatively long decision about the next steps in the (TRL 8). The definition of T development time of a propulsion development of the cryogenic reignit- Most engine components have already system, it was decided to select a two- able upper-stage engine. More precisely, reached a TRL of 5 and are close to step approach to increase Ariane-5’s in- the implementation of this objective achieving the target of TRL 6. The most the opera ydrogen environments, (including reliability, projected mass and orbit delivery capability. requires detailed studies of the engine’s significant issues to process are linked to propellant mixtur egarding ‘embrittlement’ performance, production cost, develop- The first step was the development of operating domain and assessment of its component life duration and per- feeding conditions); ment duration and cost). an adaptation of the existing Ariane-4 design through extensive, full-scale formance assessment or behaviour in – nominal perfor xperience for the engine. This exercise is initiated at an early H10 propulsion system for a new upper- engine testing in hot-firing conditions. operational conditions (pollution, exploration of stage of the launcher design because stage called ESC-A. The second step This will make it possible to increase transient conditions, etc.). The Tech- (thrust, mixture ra the FLPP Expander experience and theory show that an involved the development of an our knowledge and understanding of nological Readiness Levels of some turbopumps); oject is to overcome the overall optimised design is not necessarily adaptation of this stage to create the expander-cycle engine operations and innovative components, like the Nozzle – characterisation of ed in the assembly of optimised subsystems. ESC-B version with a new cryogenic technologies, yielding propulsion data Extension Deployment system or the stability of epare The highest performance engine, optimal engine, the ‘Vinci’. The initial ESC-B for launcher system optimisation, in igniter, are lower compared to other conditions; ble at a further stage. from a propulsion point of view, might flight was planned in 2006, following on addition to the proper development and engine technologies and require – first experience f not be the best global solution due to from the introduction of ESC-A. safeguarding of the relevant European dedicated subsystem technology tests if integrat ade-off at stage and specifications and constraints at a higher However, although Ariane-5 ECA competencies in cryogenic propulsion. It they are to be improved. system level. There is a continuous entered operational service, a should be noted that the development of There are many more significant – ated exchange of data between the engine and combination of factors including a the Vinci engine, before it was inter- technological steps in the FLPP d the launcher system throughout the downturn in the commercial market rupted, had just reached a stage that Expander Demonstrator Project that entire preliminary design process, starting delayed and then stopped the allowed the Expander Demonstrator contribute to the improvement of the – or with high-level performance data, such as development of the ESC-B stage and the Project to gain direct experience at the engine TRL. These are: the following: Vinci engine. engine hot-firing test level. – definition of an optimised starting . – thrust level; At the same time, launch system and shutdown sequence with respect – mixture ratio; studies within FLPP showed clearly the Assessment of operating domain and to progressivity, duration and – – specific impulse; need for a versatile, high-performance, design maturation propellant consumption; – restart capability; evolved cryogenic upper-stage engine The Technological Readiness Level – verification of engine stability over the – thrust-to-weight ratio; capable of delivering payloads to all – – physical interfaces; kinds of orbits, ranging from Low Earth Orbit up to exploration missions in deep Expander closed-cycle engine: how does it work? space. A high-performance upper-stage The expander thermodynamic cycle is a ‘closed cycle’, meaning that the propellants flow together engine appeared to be a central element through the thrust chamber, hence maximising the for the future launcher scenarios of the specific impulse, an indicator of engine FLPP, and a cryogenic expander engine performance. The combustion chamber pressure offered high expectations in terms of is higher than the tank pressure (~60 bars performance and reliability. compared to 2–5 bars). This pressure rise is It became quickly obvious that the ensured via two centrifugal turbopumps driven by availability of a set of expander-cycle turbines installed on the pump shafts. upper-stage engines offered a unique opportunity to progress in the The turbines are activated by the flow of high- preparation of upper-stage engines for pressure gaseous hydrogen obtained by all future launcher configurations. circulating the hydrogen pump discharge flow It was, therefore, decided at the end of around the hot combustion chamber walls. After 2005 to transfer the management and being heated up in the combustion chamber jacket, the hydrogen flows existing assets of the former Vinci through the turbines and is injected into the combustion chamber. Then, mixed with the liquid oxygen flow, it combusts and produces the hot-gas flow development to the FLPP in order to that provides the rocket engine’s thrust. form the basis of a demonstration 36 esa bulletin 134 - may 2008 www.esa.int www.esa.int esa bulletin 134 - may 2008 37 FLPP_134.qxd 6/5/08 12:37 PM Page 36 Launchers Expander Demonstrator In 1998 ESA, CNES and Arianespace project called the ‘Cryogenic Reignitable (TRL) of the engine and its decided to develop an enhanced Upper Stage Engine – Expander components, currently estimated cryogenic upper-stage for the Ariane-5 Demonstrator’. between 3 and 5, depending on the launcher in order to respond to the subjects under study, must be raised to rapid evolution of the global Project Objectives level 6 (prototype test in a relevant commercial market towards more heavy The main aim of the FLPP Expander environment), in order to properly assess payloads. In recognition of the quick Demonstrator Project is to supply the risks, cost and
Recommended publications
  • The SKYLON Spaceplane

    The SKYLON Spaceplane

    The SKYLON Spaceplane Borg K.⇤ and Matula E.⇤ University of Colorado, Boulder, CO, 80309, USA This report outlines the major technical aspects of the SKYLON spaceplane as a final project for the ASEN 5053 class. The SKYLON spaceplane is designed as a single stage to orbit vehicle capable of lifting 15 mT to LEO from a 5.5 km runway and returning to land at the same location. It is powered by a unique engine design that combines an air- breathing and rocket mode into a single engine. This is achieved through the use of a novel lightweight heat exchanger that has been demonstrated on a reduced scale. The program has received funding from the UK government and ESA to build a full scale prototype of the engine as it’s next step. The project is technically feasible but will need to overcome some manufacturing issues and high start-up costs. This report is not intended for publication or commercial use. Nomenclature SSTO Single Stage To Orbit REL Reaction Engines Ltd UK United Kingdom LEO Low Earth Orbit SABRE Synergetic Air-Breathing Rocket Engine SOMA SKYLON Orbital Maneuvering Assembly HOTOL Horizontal Take-O↵and Landing NASP National Aerospace Program GT OW Gross Take-O↵Weight MECO Main Engine Cut-O↵ LACE Liquid Air Cooled Engine RCS Reaction Control System MLI Multi-Layer Insulation mT Tonne I. Introduction The SKYLON spaceplane is a single stage to orbit concept vehicle being developed by Reaction Engines Ltd in the United Kingdom. It is designed to take o↵and land on a runway delivering 15 mT of payload into LEO, in the current D-1 configuration.
  • Rocket Propulsion Fundamentals 2

    Rocket Propulsion Fundamentals 2

    https://ntrs.nasa.gov/search.jsp?R=20140002716 2019-08-29T14:36:45+00:00Z Liquid Propulsion Systems – Evolution & Advancements Launch Vehicle Propulsion & Systems LPTC Liquid Propulsion Technical Committee Rick Ballard Liquid Engine Systems Lead SLS Liquid Engines Office NASA / MSFC All rights reserved. No part of this publication may be reproduced, distributed, or transmitted, unless for course participation and to a paid course student, in any form or by any means, or stored in a database or retrieval system, without the prior written permission of AIAA and/or course instructor. Contact the American Institute of Aeronautics and Astronautics, Professional Development Program, Suite 500, 1801 Alexander Bell Drive, Reston, VA 20191-4344 Modules 1. Rocket Propulsion Fundamentals 2. LRE Applications 3. Liquid Propellants 4. Engine Power Cycles 5. Engine Components Module 1: Rocket Propulsion TOPICS Fundamentals • Thrust • Specific Impulse • Mixture Ratio • Isp vs. MR • Density vs. Isp • Propellant Mass vs. Volume Warning: Contents deal with math, • Area Ratio physics and thermodynamics. Be afraid…be very afraid… Terms A Area a Acceleration F Force (thrust) g Gravity constant (32.2 ft/sec2) I Impulse m Mass P Pressure Subscripts t Time a Ambient T Temperature c Chamber e Exit V Velocity o Initial state r Reaction ∆ Delta / Difference s Stagnation sp Specific ε Area Ratio t Throat or Total γ Ratio of specific heats Thrust (1/3) Rocket thrust can be explained using Newton’s 2nd and 3rd laws of motion. 2nd Law: a force applied to a body is equal to the mass of the body and its acceleration in the direction of the force.
  • IAF Space Propulsion Symposium 2019

    IAF Space Propulsion Symposium 2019

    IAF Space Propulsion Symposium 2019 Held at the 70th International Astronautical Congress (IAC 2019) Washington, DC, USA 21 -25 October 2019 Volume 1 of 2 ISBN: 978-1-7138-1491-7 Printed from e-media with permission by: Curran Associates, Inc. 57 Morehouse Lane Red Hook, NY 12571 Some format issues inherent in the e-media version may also appear in this print version. Copyright© (2019) by International Astronautical Federation All rights reserved. Printed with permission by Curran Associates, Inc. (2020) For permission requests, please contact International Astronautical Federation at the address below. International Astronautical Federation 100 Avenue de Suffren 75015 Paris France Phone: +33 1 45 67 42 60 Fax: +33 1 42 73 21 20 www.iafastro.org Additional copies of this publication are available from: Curran Associates, Inc. 57 Morehouse Lane Red Hook, NY 12571 USA Phone: 845-758-0400 Fax: 845-758-2633 Email: [email protected] Web: www.proceedings.com TABLE OF CONTENTS VOLUME 1 PROPULSION SYSTEM (1) BLUE WHALE 1: A NEW DESIGN APPROACH FOR TURBOPUMPS AND FEED SYSTEM ELEMENTS ON SOUTH KOREAN MICRO LAUNCHERS ............................................................................ 1 Dongyoon Shin KEYNOTE: PROMETHEUS: PRECURSOR OF LOW-COST ROCKET ENGINE ......................................... 2 Jérôme Breteau ASSESSMENT OF MON-25/MMH PROPELLANT SYSTEM FOR DEEP-SPACE ENGINES ...................... 3 Huu Trinh 60 YEARS DLR LAMPOLDSHAUSEN – THE EUROPEAN RESEARCH AND TEST SITE FOR CHEMICAL SPACE PROPULSION SYSTEMS ....................................................................................... 9 Anja Frank, Marius Wilhelm, Stefan Schlechtriem FIRING TESTS OF LE-9 DEVELOPMENT ENGINE FOR H3 LAUNCH VEHICLE ................................... 24 Takenori Maeda, Takashi Tamura, Tadaoki Onga, Teiu Kobayashi, Koichi Okita DEVELOPMENT STATUS OF BOOSTER STAGE LIQUID ROCKET ENGINE OF KSLV-II PROGRAM .......................................................................................................................................................
  • Variations of Solid Rocket Motor Preliminary Design for Small TSTO Launcher

    Variations of Solid Rocket Motor Preliminary Design for Small TSTO Launcher

    View metadata, citation and similar papers at core.ac.uk brought to you by CORE provided by Institute of Transport Research:Publications Space Propulsion 2012 – ID 2394102 Variations of Solid Rocket Motor Preliminary Design for Small TSTO launcher Etienne Dumont Space Launcher Systems Analysis (SART), DLR, Bremen, Germany [email protected] NGL New/Next Generation Launcher Abstract SI Structural Index (mdry / mpropellant) Several combinations of solid rocket motors and ignition SRM Solid Rocket Motor strategies have been considered for a small Two Stage to TSTO Two Stage To Orbit Orbit (TSTO) launch vehicle based on a big solid rocket US Upper Stage motor first stage and cryogenic upper stage propelled by VENUS Vega New Upper Stage the Vinci engine. In order to reach the target payload avg average during the flight performance of about 1400 kg into GTO for the clean s.l. sea level version and 2700 to 3000 kg for the boosted version, the vac vacuum influence of the selected solid rocket motors on the upper 2 + 2 P23 4 P23: two ignited on ground and two with a stage structure has been studied. Preliminary structural delayed ignition designs have been performed and the thrust histories of the solid rocket motor have been tweaked to limit the upper stage structural mass. First stage and booster 1. Introduction combinations with acceptable general loads are proposed. Solid rocket motors (SRM) are commonly used for boosters or launcher first stage. Indeed they can provide high thrust levels while being compact, light and Nomenclature relatively simple compared to a liquid rocket engine Isp specific impulse s providing the same thrust level.
  • Fuel and Oxidizer Feed Systems

    Fuel and Oxidizer Feed Systems

    Fuel and Oxidizer Feed Systems Zachary Hein, Den Donahou, Andrew Doornink, Mack Bailey, John Fieler 1 1 Design Selection Recap Fuel Selection Fuel: Ethanol C2H5OH -Potential Biofuel -Low mixture ratio with LOX -Good specific impulse -Easy to get Oxidizer: Liquid Oxygen LOX -Smaller tank needed (Compared to gaseous O2) -Can be pressurized -Lowest oxidizer mixture ratio -Provides Highest specific impulse 2 Design Selection Recap Thrust Chamber Thrust Chamber Selections ● Injector: Like Impinging Doublet ● Cooling System: Regenerative Cooling ● Thrust Chamber Material: Haynes 230 3 Design Selection Recap Thrust Chamber Thrust Chamber Selections ● Injector: Like Impinging Doublet ● Cooling System: Regenerative Cooling ● Thrust Chamber Material: Haynes 230 Huzel, Dieter, and David Huang. "Introduction." Modern Engineering for Design of Liquid-Propellant Rocket Engines. Vol. 147. Washington D.C.: AIAA, 1992. 7-22. Print. 4 Design Selection Recap Thrust Chamber Thrust Chamber Selections ● Injector: Like Impinging Doublet ● Cooling System: Regenerative Cooling ● Thrust Chamber Material: Haynes 230 Huzel, Dieter, and David Huang. "Introduction." Modern Engineering for Design of Liquid-Propellant Rocket Engines. Vol. 147. Washington D.C.: AIAA, 1992. 7-22. Print. http://www.k-makris.gr/RocketTechnology/ThrustChamber/Thrust_Chamber.htm 5 Design Selection Recap Thrust Chamber Thrust Chamber Selections ● Injector: Like Impinging Doublet ● Cooling System: Regenerative Cooling ● Thrust Chamber Material: Haynes 230 Huzel, Dieter, and David Huang. "Introduction." Modern Engineering for Design of Liquid-Propellant Rocket Engines. Vol. 147. Washington D.C.: AIAA, 1992. 7-22. Print. http://www.k-makris.gr/RocketTechnology/ThrustChamber/Thrust_Chamber.htm http://www.alibaba.com/product-detail/haynes-seamless-pipe_1715659362.html 6 Turbo Pump Basics Turbo Pumps provide pressurization to gaseous fuel components to required pressures and mixture ratios.
  • Basic Analysis of a LOX/Methane Expander Bleed Engine

    Basic Analysis of a LOX/Methane Expander Bleed Engine

    DOI: 10.13009/EUCASS2017-332 7TH EUROPEAN CONFERENCE FOR AERONAUTICS AND AEROSPACE SCIENCES (EUCASS) DOI: ADD DOINUMBER HERE Basic Analysis of a LOX/Methane Expander Bleed Engine ? ? ? Marco Leonardi , Francesco Nasuti † and Marcello Onofri ?Sapienza University of Rome Via Eudossiana 18, Rome, Italy [email protected] [email protected] [email protected] · · †Corresponding author Abstract As present trends in rocket engine development recommend overall simplicity and reliability as the main design driver, while preserving high performance, expander cycle engines based on the oxygen-methane pair have been considered as a possible upper stage option. A closed expander cycle is considered for Vega Evolution upper stage, while there are no studies published in the literature on methane-based expander bleed cycles. A basic cycle analysis is presented to evaluate the performance of an oxygen/methane ex- pander bleed cycle for an engine of 100 kN thrust class. Results show the feasibility of the system and its peculiarities with respect to the better known expander bleed cycle based on hydrogen. 1. Introduction The high chamber pressure required to achieve high specific impulse in liquid propellant rocket engines (LRE), has been efficiently obtained by pump-fed systems. Different solutions have been proposed since the beginning of space age and just a few of them has found its own field of application. In these systems the pumps are driven by gas turbines whose power comes from two possible sources: combustion or cooling system. The different needs for the specific applications (booster, sustainer or upper stage of different classes of rockets) led to classify pump-fed LRE systems in open and closed cycles, which differ because of turbine discharge pressure.14, 16 Closed cycles are those providing the best performance because the whole propellant mass flow rate is exploited in the main chamber.
  • Experimental Study of the Combustion Efficiency in Multi-Element Gas

    Experimental Study of the Combustion Efficiency in Multi-Element Gas

    energies Article Experimental Study of the Combustion Efficiency in Multi-Element Gas-Centered Swirl Coaxial Injectors Seongphil Woo 1,* , Jungho Lee 1,2, Yeoungmin Han 2 and Youngbin Yoon 1,3 1 Department of Aerospace Engineering, Seoul National University, Seoul 08826, Korea; [email protected] (J.L.); [email protected] (Y.Y.) 2 Korea Aerospace Research Institute, Daejeon 34133, Korea; [email protected] 3 Institute of Advanced Aerospace Technology, Seoul National University, Seoul 08826, Korea * Correspondence: [email protected] Received: 27 October 2020; Accepted: 16 November 2020; Published: 19 November 2020 Abstract: The effects of the momentum-flux ratio of propellant upon the combustion efficiency of a gas-centered-swirl-coaxial (GCSC) injector used in the combustion chamber of a full-scale 9-tonf staged-combustion-cycle engine were studied experimentally. In the combustion experiment, liquid oxygen was used as an oxidizer, and kerosene was used as fuel. The liquid oxygen and kerosene burned in the preburner drive the turbine of the turbopump under the oxidizer-rich hot-gas condition before flowing into the GCSC injector of the combustion chamber. The oxidizer-rich hot gas is mixed with liquid kerosene passed through combustion chamber’s cooling channel at the injector outlet. This mixture has a dimensionless momentum-flux ratio that depends upon the dispensing speed of the two fluids. Combustion tests were performed under varying mixture ratios and combustion pressures for different injector shapes and numbers of injectors, and the characteristic velocities and performance efficiencies of the combustion were compared. It was found that, for 61 gas-centered swirl-coaxial injectors, as the moment flux ratio increased from 9 to 23, the combustion-characteristic velocity increased linearly and the performance efficiency increased from 0.904 to 0.938.
  • Jarvis Heavy Launch Vehicle

    Jarvis Heavy Launch Vehicle

    JARVIS HEAVY LAUNCH VEHICLE By Forum Orbiter Italia Version 2.62 – October 2012 USER MANUAL Disclaimer and credits This add-on is provided “as is”, without any kind of warranty; it is compatible with Orbiter 2006-P1 (build 060929) and with Orbiter 2010-P1 (build 100830). Many thanks to Dr. Martin Schweiger, for the Orbiter Space Simulator. For the others developers: You are free to use parts of our work, eg sound and texture, but you must credit us as the original source of your work. FOI Credits - Andrew: add-on conception; rocket textures, meshes and configuration; documentation editing. - Fausto: new launch pad textures, meshes and configuration. - Pete Conrad: engine meshes and textures; Shuttle SRB meshes and textures; “dummy” payload meshes and textures. - FedeX: beta testing. - Dany: “Forum Orbiter Italia” logo. - Ripley: D3D9/D3D11 documentation. Forum Orbiter Italia: http://orbiteritalia.forumotion.com/ Introduction In the mid-eighties, the "Jarvis" project was the last serious attempt to revive the glorious Saturn V rocket, and at the same time, one of the first ideas of an alternative use for the Space Shuttle hardware, many years before the current "Ares", "Direct" and “SLS” projects. The Jarvis rocket combines the powerful Apollo-era F-1 and J-2 engines with Space Shuttle electronics and 8.4 m stages (the same size of the Shuttle External Tank). Later versions, with Space Shuttle Main Engines (SSME) and/or Solid Rocket Boosters (SRB), were proposed, but never realized. Forum Orbiter Italia has developed a complete and versatile family of heavy launchers around these original ideas and projects.
  • A Survey of Automatic Control Methods for Liquid-Propellant Rocket Engines

    A Survey of Automatic Control Methods for Liquid-Propellant Rocket Engines

    A survey of automatic control methods for liquid-propellant rocket engines Sergio Pérez-Roca, Julien Marzat, Hélène Piet-Lahanier, Nicolas Langlois, Francois Farago, Marco Galeotta, Serge Le Gonidec To cite this version: Sergio Pérez-Roca, Julien Marzat, Hélène Piet-Lahanier, Nicolas Langlois, Francois Farago, et al.. A survey of automatic control methods for liquid-propellant rocket engines. Progress in Aerospace Sciences, Elsevier, 2019, pp.1-22. 10.1016/j.paerosci.2019.03.002. hal-02097829 HAL Id: hal-02097829 https://hal.archives-ouvertes.fr/hal-02097829 Submitted on 12 Apr 2019 HAL is a multi-disciplinary open access L’archive ouverte pluridisciplinaire HAL, est archive for the deposit and dissemination of sci- destinée au dépôt et à la diffusion de documents entific research documents, whether they are pub- scientifiques de niveau recherche, publiés ou non, lished or not. The documents may come from émanant des établissements d’enseignement et de teaching and research institutions in France or recherche français ou étrangers, des laboratoires abroad, or from public or private research centers. publics ou privés. A survey of automatic control methods for liquid-propellant rocket engines Sergio Perez-Roca´ a,c,∗, Julien Marzata,Hel´ ene` Piet-Lahaniera, Nicolas Langloisb, Franc¸ois Faragoc, Marco Galeottac, Serge Le Gonidecd aDTIS, ONERA, Universit´eParis-Saclay, Chemin de la Huniere, 91123 Palaiseau, France bNormandie Universit´e,UNIROUEN, ESIGELEC, IRSEEM, Rouen, France cCNES - Direction des Lanceurs, 52 Rue Jacques Hillairet, 75612 Paris, France dArianeGroup SAS, Forˆetde Vernon, 27208 Vernon, France Abstract The main purpose of this survey paper is to review the field of convergence between the liquid-propellant rocket- propulsion and automatic-control disciplines.
  • Materials for Liquid Propulsion Systems

    Materials for Liquid Propulsion Systems

    CHAPTER 12 Materials for Liquid Propulsion Systems John A. Halchak Consultant, Los Angeles, California James L. Cannon NASA Marshall Space Flight Center, Huntsville, Alabama Corey Brown Aerojet-Rocketdyne, West Palm Beach, Florida 12.1 Introduction Earth to orbit launch vehicles are propelled by rocket engines and motors, both liquid and solid. This chapter will discuss liquid engines. The heart of a launch vehicle is its engine. The remainder of the vehicle (with the notable exceptions of the payload and guidance system) is an aero structure to support the propellant tanks which provide the fuel and oxidizer to feed the engine or engines. The basic principle behind a rocket engine is straightforward. The engine is a means to convert potential thermochemical energy of one or more propellants into exhaust jet kinetic energy. Fuel and oxidizer are burned in a combustion chamber where they create hot gases under high pressure. These hot gases are allowed to expand through a nozzle. The molecules of hot gas are first constricted by the throat of the nozzle (de-Laval nozzle) which forces them to accelerate; then as the nozzle flares outwards, they expand and further accelerate. It is the mass of the combustion gases times their velocity, reacting against the walls of the combustion chamber and nozzle, which produce thrust according to Newton’s third law: for every action there is an equal and opposite reaction. [1] Solid rocket motors are cheaper to manufacture and offer good values for their cost. Liquid propellant engines offer higher performance, that is, they deliver greater thrust per unit weight of propellant burned.
  • Development of LE-X Engine,Mitsubishi Heavy Industries

    Development of LE-X Engine,Mitsubishi Heavy Industries

    Mitsubishi Heavy Industries Technical Review Vol. 48 No. 4 (December 2011) 36 Development of the LE-X Engine MASAHIRO ATSUMI*1 KIMITO YOSHIKAWA*2 AKIRA OGAWARA*3 TADAOKI ONGA*3 The expander bleed cycle is an engine cycle that was developed in Japan for practical applications. It has robust operational characteristics against disturbances due to its simplicity, and was adopted for the LE-5B engine, the second-stage engine of H-IIA launch vehicle. The LE-5B has many capabilities, providing restart capability, throttling and idle mode combustion (extremely low thrust operations) capability, these capabilities is evaluated highly in the world due to its reliable performance. The expander bleed cycle was first adapted for the LE-5A engine, which was an improved version of the LE-5 engine. The LE-5 was the first Japanese liquid oxidant/hydrogen (LOX/LH2) engine. This cycle was also adapted for the LE-5B engine, and more recently, for the MB-XX engine, which is a cooperative development between Mitsubishi Heavy Industries, Ltd. (MHI) and U.S. Pratt and Whitney Rocketdyne. These are all second-stage engines. Now, MHI is adopting this cycle for the first-stage engine of a next-generation launch vehicle under contract with the Japan Aerospace Exploration Agency (JAXA), with the intention of providing world-standard first-class reliability. This report describes the features of the expander bleed engine cycle and our approach for providing the highly reliable LE-X engine. |1. Introduction In 1999, H-II Launch Vehicle Flight No. 8 ended in failure due to an explosion in the first-stage LE-7 engine during flight.
  • Lox/Lch4 Upper Stage Development Strategies for Future Launchers

    Lox/Lch4 Upper Stage Development Strategies for Future Launchers

    TURBO, vol. VI (2019), no. 2 LOX/LCH4 UPPER STAGE DEVELOPMENT STRATEGIES FOR FUTURE LAUNCHERS Theodora ANDREESCU1, Andreea MANGRA1, Valeriu VILAG1, Ion MALAEL1, Alexandru CANCESC1, Jeni VILAG1, Dan IFRIM1, Simona DANESCU1 ABSTRACT: The reduction of Earth-to-orbit launch costs in conjunction with an increase in Launcher reliability and operational efficiency are the key requirements of future space transportation systems. This paper underlines the progress in LOX/CH4 upper stage engine development carried out by COMOTI and also being provided the prediction of the rocket engine performances at the conceptual and preliminary stages of design. This paper focuses on the trade-off studies for the engine architecture definition, considering both open and closed thermodynamic cycles. Various subsystems configurations have been taken into account, analyzing the optimum configuration in terms of performance. The main operating and geometrical parameters were discussed: combustion pressure, optimum mixture ratio, turbine pressure ratio, thrust chamber geometry, and the turbopump size is addressed. KEYWORDS: upper stage, liquid rocket engine, turbopump, LOX/CH4 cryogenic propellant, thrust chamber NOMENCLATURE A – turbine characteristic area 풎̇ – gas flow rate 푨풆 - nozzle exit diameter 푷풑풐풙 – liquid oxygen pump power 푨풕 – nozzle throat diameter 푷풑풇풖풆풍 – liquid methane pump power 푨풄 – combustion chamber area 푷푻 – turbine power 푻풄 – combustion temperature 풏풔 – pump specific rotational speed 휸 - specific heat ratio 휶 – turbine flow angle ∗ 푪풑- specific heat capacity 풑 - turbine pressure losses 풑풄 – combustion pressure R – gas constant 푪푭 – thrust coefficient – density 풄∗ - characteristic velocity M – Mach number 품ퟎ – gravitational constant 풉풑풇, 풉풑풎 – turbine blades hight 풙풑풇, 풙풑풎, - axial width 1. INTRODUCTION As the present trend in rocket engine development recommends a high versatility and low launch service cost, while preserving high performance, expander cycle upper stage based on LOX/LCH4 being a key competitiveness factor recognized by the market.